Turbine airfoil for gas turbine engine

ABSTRACT

A hollow cooled turbine airfoil is provided having pressure and suction side walls and a plurality of trailing edge cooling passages that feed cooling air bleed slots at the trailing edge. The suction side wall has a selectively reduced thickness, which allows shortened trailing edge slots, improving trailing edge cooling.

BACKGROUND OF THE INVENTION

[0001] The present invention relates generally to gas turbine engines,and more particularly to hollow air cooled airfoils used in suchengines.

[0002] A gas turbine engine includes a compressor that providespressurized air to a combustor wherein the air is mixed with fuel andignited for generating hot combustion gases. These gases flow downstreamto one or more turbines that extract energy therefrom to power thecompressor and provide useful work such as powering an aircraft inflight. In a turbofan engine, which typically includes a fan placed atthe front of the core engine, a high pressure turbine powers thecompressor of the core engine. A low pressure turbine is disposeddownstream from the high pressure turbine for powering the fan. Eachturbine stage commonly includes a stationary turbine nozzle followed inturn by a turbine rotor.

[0003] The turbine rotor comprises a row of rotor blades mounted to theperimeter of a rotor disk that rotates about the centerline axis of theengine. Each rotor blade typically includes a shank portion having adovetail for mounting the blade to the rotor disk and an airfoil thatextracts useful work from the hot gases exiting the combustor. A bladeplatform, formed at the junction of the airfoil and the shank portion,defines the radially inner boundary for the hot gas stream. The turbinenozzles are usually segmented around the circumference thereof toaccommodate thermal expansion. Each nozzle segment has one or morenozzle vanes disposed between inner and outer bands for channeling thehot gas stream into the turbine rotor.

[0004] The high pressure turbine components are exposed to extremelyhigh temperature combustion gases. Thus, the turbine blades and nozzlevanes typically employ internal cooling to keep their temperatureswithin certain design limits. The airfoil of a turbine rotor blade, forexample, is ordinarily cooled by passing cooling air through an internalcircuit. The cooling air normally enters through a passage in theblade's root and exits through film cooling holes formed in the airfoilsurface, thereby producing a thin layer or film of cooling air thatprotects the airfoil from the hot gases. Known cooling arrangementsoften include a plurality of openings in the trailing edge through whichcooling air is discharged. These openings may take the form of holes, orof a pressure side bleed slot arrangement, in which the airfoil pressureside wall stops short of the extreme trailing edge of the airfoil,creating an opening which is divided into individual bleed slots by aplurality of longitudinally extending lands incorporated into theairfoil casting. These slots perform the function of channeling a thinfilm of cooling air over the surface of the airfoil trailing edge.Airfoils having such a pressure side bleed slot arrangement are known tobe particularly useful for incorporating a thin trailing edge. Ineffect, the trailing edge thickness of the airfoil is equal to that ofthe suction side thickness alone. This is desirable in terms ofaerodynamic efficiency. However, a very thin trailing edge typicallyresults in a relatively long bleed slot, which reduces coolingeffectiveness, because of mixing of the hot combustion gases flowingover the exterior of the blade with the cooling air flow passing throughthe slots. Accordingly, there is a need for improved cooling of airfoiltrailing edges while maintaining the aerodynamic efficiency thereof.

BRIEF SUMMARY OF THE INVENTION

[0005] The above-mentioned need is met by the present invention, whichprovides a turbine airfoil having pressure and suction side walls and aplurality of trailing edge cooling passages that feed cooling air bleedslots at the trailing edge. The suction side wall has a varyingthickness. The minimum thickness portion is positioned so as to allowshortened trailing edge slots, thereby improving trailing edge cooling.

[0006] The present invention and its advantages over the prior art willbecome apparent upon reading the following detailed description and theappended claims with reference to the accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

[0007] The subject matter that is regarded as the invention isparticularly pointed out and distinctly claimed in the concluding partof the specification. The invention, however, may be best understood byreference to the following description taken in conjunction with theaccompanying drawing figures in which:

[0008]FIG. 1 is a perspective view of a turbine blade embodying thecooling configuration of the present invention.

[0009]FIG. 2 is a partial cross-sectional view of a turbine blade takenalong line 2-2 of FIG. 1.

[0010]FIG. 3 is a partial cross-sectional view of a prior art turbineblade.

DETAILED DESCRIPTION OF THE INVENTION

[0011] Referring to the drawings wherein identical reference numeralsdenote the same elements throughout the various views, FIG. 1illustrates an exemplary turbine blade 10. The turbine blade 10 includesa conventional dovetail 12, which may have any suitable form includingtangs that engage complementary tangs of a dovetail slot in a rotor disk(not shown) for radially retaining the blade 10 to the disk as itrotates during operation. A blade shank 14 extends radially upwardlyfrom the dovetail 12 and terminates in a platform 16 that projectslaterally outwardly from and surrounds the shank 14. A hollow airfoil 18extends radially outwardly from the platform 16 and into the hot gasstream. The airfoil 18 has a concave pressure side wall 20 and a convexsuction side wall 22 joined together at a leading edge 24 and at atrailing edge 26. The airfoil 18 may take any configuration suitable forextracting energy from the hot gas stream and causing rotation of therotor disk. The blade 10 is preferably formed as a one-piece casting ofa suitable superalloy, such as a nickel-based superalloy, which hasacceptable strength at the elevated temperatures of operation in a gasturbine engine. The blade incorporates a number of trailing edge bleedslots 28 on the pressure side 20 of the airfoil. The bleed slots 28 areseparated by a number of longitudinally extending lands 30. At least aportion of the airfoil is typically coated with a protective coating 31(FIG. 2), such as an environmentally resistant coating, or a thermalbarrier coating, or both.

[0012]FIG. 3 illustrates a typical prior art arrangement of a portion ofthe trailing edge of a turbine blade. Pressure side wall 52 and suctionside wall 50 are separated by an internal cavity 54. The side wallstaper inwards toward the trailing edge 60. The suction side wallcontinues unbroken the entire length of the blade all the way to thetrailing edge 60, whereas the pressure side wall 52 has an aft-facinglip 66 so as to expose an opening in the trailing edge 60, which isdivided by lands 57 into a plurality of trailing edge slots 58. Theaft-facing lip 66 defines the position of the trailing edge coolingpassage exit 62. In this type of turbine blade, the trailing edgethickness at the aft end of the blade is essentially equal to thethickness of the suction side wall 50 alone, as described above.

[0013] The cooling effectiveness of the trailing edge slots 58 isrelated to their length L, which is the distance from the trailing edgecooling passage exit 62 to the trailing edge 60. This dimension is alsosometimes referred to as the slot breakout distance. The longer the slotlength L, the less is the trailing edge cooling effectiveness becausethe hot flowpath gases passing over the airfoil upstream of the extremetrailing edge tend to mix with the cooling air discharged from thetrailing edge cooling passages 56. Effective cooling of the trailingedge 60 is especially important in high pressure turbine airfoilapplications for life, durability, and reparability.

[0014] The trailing edge slot length L is controlled by severalvariables. The wedge angle W is the included angle between the outersurfaces of the airfoil and is typically measured towards the aft end ofthe airfoil, where the airfoil surfaces have the least curvature. Thetrailing edge thickness t is defined as the wall thickness at apredetermined small distance, for example 0.762 mm (0.030 in.), from theextreme aft end of the airfoil. The combination of the wedge angle W andthe trailing edge thickness t determine the maximum overall airfoilthickness at each location along the aft portion of the airfoil. Theoverall airfoil thickness at the exit 62 of the trailing edge coolingpassage 56 is denoted T and has a certain minimum dimension, asdescribed more fully below. It would be possible to decrease the slotlength L by increasing the wedge angle W, thus increasing dimension T.However, increasing wedge angle W and therefore the overall airfoilthickness would have a detrimental effect on aerodynamic performance.Dimension T is the sum of the pressure side wall thickness P, thesuction side wall thickness S, and the trailing edge cooling passagewidth H. Reduction of dimensions P, S, or H would allow the slot lengthL to be reduced without increasing dimension T. However, there is aminimum trailing edge hole width H required in order to avoid excessivebreakage of the ceramic cores used to produce the passages 56 during thecasting process of the blade 10 and to provide the required coolingairflow. Also, there is a minimum thickness P required of the pressureside wall 52 and a minimum thickness S required of the suction side wall50 for mechanical integrity.

[0015] The present invention, illustrated in FIG. 2, avoids thesedifficulties by selectively decreasing the suction side wall thickness.In the illustrated airfoil, an internal cavity 34 is bounded by pressureside wall 20 and suction side wall 22. The pressure side wall 20 has anouter or “hot” side 40 exposed to the flow of combustion gases and aninner or “cold” side 42. Likewise, the suction side wall has an outer or“hot” side 36 exposed to the flow of combustion gases and an inner or“cold” side 38. A trailing edge cooling passage 32 connects the internalcavity 34 with the trailing edge bleed slot 28 and is bounded by thecold side 42 of the pressure side wall 20 and the cold side 38 of thesuction side wall 22. The trailing edge cooling passage is bounded inthe radial direction by lands 30 (only one of which is shown in FIG. 2).The pressure side wall 20 has an aft-facing lip 48 defining the forwardend of slot 28. In the exemplary embodiment shown, the wedge angle WW,pressure side wall thickness D, trailing edge passage width E, trailingedge thickness C, and the blade overall thickness TT at the trailingedge cooling passage exit 44 are unchanged relative to a nominal orbaseline design. The thickness of the suction side wall 22 is smoothlyvaried so that the thickness near the trailing edge cooling passage exit44 is less than the thickness at the inlet of the passage 32 and thetrailing edge thickness C. In varying the suction side wall thickness,the hot side 36 of suction side wall 22 is not changed relative to thebaseline design. Rather, the contour of the wall cold side 38 ismodified. In this way, the external contour definition and thus theaerodynamic performance of the blade are unchanged from a baselineairfoil.

[0016] More specifically, The thickness A of the suction side wall 22 ata point upstream of the trailing edge cooling hole exit is a nominalthickness, which is determined taking into consideration the expectedoperating environment, including thermal, aerodynamic, and mechanicalloads, and production process capability, including the possibility of“core shift” which is an effect caused by the movement of the ceramiccores used to define the interior dimensions of the blade during thecasting process. Core shift can produce an unacceptability thin wall ifallowances are not made for the effect. Within these constraints, thewall thickness is made as small as possible to minimize the materialused, and thus the weight of the blade. A representative minimum valueof dimension A may be about 0.737 mm (0.029 in.). The thickness ofsuction side wall 22 is smoothly tapered in a curved shape as it extendstowards the trailing edge 26, resulting in a reduced thickness B at theaxial location of the exit 44 of trailing edge cooling passage 32. Thisvalue is typically about 0.076-0.100 mm (0.003-0.004 in.) less thanthickness A, or about 0.635-0.660 mm (0.025-0.026 in.).

[0017] In the exemplary embodiment illustrated, the minimum thickness ofthe suction side wall 22 occurs slightly aft of dimension B because ofthe natural taper of the suction side wall 22 upstream of the exit 44and the method used to generate the wall contour. In the illustratedembodiment, the pressure side wall thickness D at the lip 48, thetrailing edge cooling passage width E at the exit 44, and the trailingedge thickness C are all fixed. The contour of the suction side wall 22is then extended rearward and subsequently a smooth curve is used togenerate the intervening portion of suction side wall 22. However, thesuction side wall minimum thickness could occur at the exit 44 orforward of the exit 44 if desired. Aft of the minimum thickness section,the thickness of suction side wall 22 is smoothly increased in the aftdirection, until at the trailing edge 26, thickness C is equal to thebaseline value dictated by the aerodynamic design (which isapproximately equal to thickness A, but may also be greater than A).Thickness C may be about 0.737 mm (0.029 in.). The incorporation of areduced thickness B of the suction side wall 22 while maintainingpressure side wall thickness D and trailing edge cooling hole width Ereduces overall airfoil thickness TT (which is the sum of dimensions B,D, and E). This allows pressure side wall 20 to be extended rearward,moving the exit 44 of trailing edge cooling passage 32 towards thetrailing edge 26 and reducing the slot length LL. The slot length LLcould not be shortened in this manner without the curve section inpressure side wall 22, as can be seen by reference to the line marked 46in FIG. 3, which represents a straight line tangent to pressure sidewall 22 at the axial locations of dimensions A and C.

[0018] In an exemplary embodiment, the reduction in pressure side wallthickness B to about 0.635-0.660 mm (0.025-0.026 in.), or about 12% lessthan thickness A, allows a reduction in slot length LL of about 1.52 mm(0.060 in.), or about 32% of the baseline slot length. This can producea significant reduction in trailing edge temperatures due to improvedcooling. Analysis of this design indicates that the reduction in slotlength by about 1.52 mm (0.060 in.) would reduce the expectedtemperature of the trailing edge by about 16° C. (30° F.), in comparisonto the baseline design with an unchanged slot length. This specifiedvalue of thickness B would cause the wall thickness of productionairfoils to approach the allowable minimum limits of in an area of theairfoil subject to appreciable core shift, for example at the locationof thickness A. However, because the core used to produce the trailingedge cooling passage 32 during the casting process is better restrainednear the trailing edge cooling passage exit 44 than at the upstreamportion of the airfoil, the thickness B may be reduced below a nominalvalue without creating a risk of having an unacceptably thin wall. Byincorporating a thin section B only where needed to accommodate theshorter slot length, the beneficial effect of a shortened slot isachieved while minimizing the risk to part integrity and productionyield.

[0019] Because of the relatively large change in slot length LLpermitted by a small change in thickness B, even a very small reductionin thickness B will enable the slot length LL to be reduced in abeneficial manner. Practically speaking, however, too small of areduction in thickness B will be smaller than the variation in thedimensions that can be reliably produced in an actual airfoil.Accordingly, about 0.076 mm (0.003 in.) represents a practical minimumreduction of thickness B. However, to the extent that technologiesbecome available to reliably produce reductions of less than this value,for example improved manufacturing processes, the benefit of the presentinvention would still be obtained, albeit to a lesser degree because ofthe lesser amount of shortening of the slot length LL. On the otherhand, greater reductions of thickness B would allow further shorteningof the slot length LL, resulting in additional temperature reductions.Reduction in the thickness B is practically limited. The thinner thewall, the greater the chances of part failure in manufacture oroperation. Although the practical minimum thickness varies dependingupon the particular airfoil and operating conditions, a representativeextreme minimum value of B is about 0.038-0.051 mm (0.015-0.020 in.). Tothe extent that technologies in manufacturing and/or materials wouldenable the reliable production of thinner airfoil walls, this wouldallow the benefit of the present invention to be obtained through evenfurther shortening of the slot length LL.

[0020] Additionally, because the trailing edge has a “club” shape causedby the curvature of the suction side wall 22 as described above, aprojected area of the wall surface 38 downstream of the lip 48 isexposed to impingement of the cooling air flow from the slot 28, inaddition to film cooling. This improves the heat transfer coefficientand reduces the temperature of the pressure side wall. The morepronounced the club shape, the greater the impingement effect.

[0021] The present invention has been described in conjunction with anexemplary embodiment of a turbine blade. However, it should be notedthat the invention is equally applicable to any hollow airfoil,specifically including airfoils for stationary turbine nozzles (orvanes) as well as rotating blades.

[0022] The foregoing has described a turbine airfoil having improvedcooling through incorporation of a shortened trailing edge slot. Whilespecific embodiments of the present invention have been described, itwill be apparent to those skilled in the art that various modificationsthereto can be made without departing from the spirit and scope of theinvention as defined in the appended claims.

What is claimed is:
 1. An airfoil having a leading edge and a trailingedge, comprising: a pressure side wall extending from said leading edgeto said trailing edge; a suction side wall extending from said leadingedge to said trailing edge; an internal cavity; a slot disposed in saidpressure side wall adjacent said trailing edge; and a passage disposedbetween said internal cavity and said slot, said passage having an inletin fluid communication with said internal cavity and an exit in fluidcommunication with said slot, said passage being bounded by saidpressure side wall and said suction side wall, wherein said suction sidewall has an inner surface and a varying thickness, wherein the value ofsaid thickness at said inlet and said trailing edge is greater than saidthickness near said exit.
 2. The airfoil of claim 1 wherein said innersurface is curved.
 3. The airfoil of claim 1 wherein said thickness isat a minimum value at said exit.
 4. The airfoil of claim 1 wherein saidthickness is at a minimum aft of said exit.
 5. The airfoil of claim 1wherein said thickness is at a minimum forward of said exit.
 6. Anairfoil having a leading edge and a trailing edge, comprising: a suctionside wall extending from said leading edge to said trailing edge; apressure side wall extending from said leading edge to said trailingedge and having an aft-facing lip near an aft end thereof; an internalcavity; a slot disposed in said pressure side wall adjacent saidtrailing edge; and a passage disposed between said internal cavity andsaid slot, said passage having an inlet in fluid communication with saidinternal cavity and an exit in fluid communication with said slot, saidpassage being bounded by said pressure side wall and said suction sidewall, wherein said suction side wall has a curved inner surface suchthat the distance between said lip and said trailing edge is minimizedfor a given thickness of said lip and a given width of said passage atsaid exit.
 7. The airfoil of claim 6 wherein the thickness of saidsuction side wall at said inlet and said trailing edge is greater thanthe thickness of said suction side wall near said exit.
 8. The airfoilof claim 6 wherein said thickness of said suction side wall is at aminimum value at said exit.
 9. The airfoil of claim 6 wherein saidthickness of said suction side wall is at a minimum aft of said exit.10. the airfoil of claim 6 wherein said thickness of said suction sidewall is at a minimum forward of said exit.